High-power ion propulsion systems have been shown to be capable of providing substantial benefits for the exploration of space. Considerable useful background material is contained in the following papers: (1) Brophy, J. R. and Barnett, J. W., "Benefits of Electric Propulsion for the Space Exploration Initiative," AIAA Paper No. 90-2756, July 1990; (2) Gilland, J. H., Myers, R. M., and Patterson, M. J., "Multimegawatt Electric Propulsion System Design Considerations," AIAA Paper No. 90-2552, July 1990; (3) Frisbee, R. H., et al., "Advanced Propulsion Options for the Mars Cargo Mission." AIAA Paper No. 90-1997, July 1990; (4) Hack, K. J., et al., "Evolutionary Use of Nuclear Electric Propulsion," AIAA Paper No. 90-3821, September 1990; (5) Galecki, D. L., and Patterson, M. J., "Nuclear Powered Mars Cargo Transport Mission Utilizing Advanced Ion Propulsion," NASA TM 100109, July 1987; (6) Gilland, J. H., "Synergistic Use of High and Low Thrust Propulsion Systems for Piloted Missions to Mars," AIAA-91-2346, June 1991; and (7) Frisbee, R. H., Blandino, J. J., and Leifer, S. D., "A Comparison of Chemical Propulsion, Nuclear Thermal Propulsion, and Multimegawatt Electric Propulsion for Mars Missions," AIAA-91-2332, June 1991.
However, for ion propulsion to fulfill its promise requires the development of ion engines which can process input powers on the order of hundreds to thousands of kilowatts at specific impulses in the range 7,000 to 10,000 seconds with useful lifetimes of 10,000 hours. From 1961 to approximately 1981 most ion engine research focused on the use of mercury as the propellant. A 150-cm diameter mercury ion engine was operated at input powers as high as 130 kW with a specific impulse of 8,150 seconds and an overall efficiency of 70%, as reported in the paper by Nakanishi, S. and Pawlik, E. V., "Experimental Investigation of a 1.5 m-diam. Kaufman Thruster," J. Spacecraft, Vol. 5, No. 7, July 1968, pp. 801-807.
In other work a mercury ion engine was operated at specific impulses greater than 16,000 seconds, as reported in the paper by Byers, D. C., "An Experimental Investigation of a High-Voltage Electron-Bombardment Ion Thruster," NASA TM X-52429, May 1968. The J-Series mercury ion thruster, which was designed for a maximum input power of 2.7 kW at a specific impulse of 3,000 seconds, was developed to nearly flight readiness for use in the Solar Electric Propulsion Stage (SEPS), as reported in the paper by Bechtel, R. T., "The 30 cm J Series Mercury Bombardment Thruster," AIAA Paper No. 81-0714, April 1981.
Since 1981 most ion propulsion research has centered on the use of noble gas propellants, with engine sizes ranging from 10 cm to 50 cm. The following papers and their references report much of that research: Fearn, D. G., "The Control Philosophy of the UK-10 and UK-25 Ion Thruster," AIAA Paper No. 90-2629, July 1990; Latham, P. M., Martin A. R., and Bond, A., "Design Manufacture and Performance of the UK-25 Engineering Model Thruster," AIAA Paper No. 90-2541, July 1990; Bassner, H., "Status of the Space Testing Programs of the RF-Ion Thruster RIT-10," AIAA Paper No. 91-1889, June 1991; Groh, K. H., et al., "Inert Gas Performance of the RIT 35 Main Propulsion Unit," IEPC-88-098, presented at the 20th International Electric Propulsion Conference, Garmisch-Partenkirchen, Germany, October 1988; Beattie, J. R. and Matossian, J. N., "Xenon Ion Propulsion for Stationkeeping and Orbit Rasing, "IEPC-88-052, presented at the 20th International Electric Propulsion Conference, Oct. 1988; Patterson, M. J. and Verhey, T. R., "5 kW Xenon Ion Thruster Life Test," AIAA Paper No. 90-2543, July 1990; Patterson, M. J. and Rawlin, V. K., "Performance of 10-kW Class Xenon Ion Thrusters," AIAA 88-2914, July 1988; Nakamura, Y., Matsumoto, M., Kitamura, S., and Miyazaki, K., "Discharge Performance of a 12 cm Cusp Xenon Ion Thruster," IEPC 88-061, presented at the 20th International Electric Propulsion Conference in Garmisch-Parternkirchen, Germany, October 1988; and Yamagiwa, Y., et al., "A 30-cm Diameter Xenon Ion Thruster--Design and Initial Test Results," IEPC 88-095, presented at the 20th International Electric Propulsion Conference in Garmisch-Partenkirchen, Germany, October 1988.
The 30-cm diameter J-Series thruster has been operated at input powers up to 17 kW with a specific impulse of 4,400 seconds using xenon propellant, as reported by Patterson, M. J. and Rawlin, V. K., "Performance of 10-kW Class Xenon Ion Thrusters," AIAA 88-2914, July 1988. The same paper also reports a 50-cm diameter thruster has been operated at up to 20 kW at a specific impulse of 4,600 seconds, again with xenon.
Ion engines operate by ionizing the propellant gas through electron bombardment and then accelerating the resulting positive ions electrostatically. The magnitude of the applied high voltage which accelerates the ions and the ion charge-to-mass ratio determines the exhaust velocity. Typically greater than 85% of the input power is processed by the positive high-voltage supply which accelerates the ions. Most of the remaining 15% of the input power goes into creating the ions and is supplied by a separate discharge power supply as indicated in the generic power supply schematic shown in FIG. 1.
The attractive feature of ion propulsion is that the electrostatic acceleration process is almost 100% efficient. In practice the acceleration efficiency is typically 99.7% This nearly lossless acceleration mechanism enables the development of ion engines which can process megawatts of input power while maintaining reasonable engine component temperatures without active cooling. It also is responsible for the high overall engine efficiencies characteristic of ion propulsion. Furthermore, this feature guarantees that scaling ion engines up to megawatt power levels is rewarded with an engine efficiency close to that of prior-art efforts.
Space charge effects in the accelerator system of ion engines place an upper limit on the thrust density (and hence power density) which ion engines can achieve at a given specific impulse. Therefore, to increase the power and thrust capabilities of an ion engine it is necessary to increase the area of the ion accelerator system while maintaining a constant thrust density.
For conventional ion engines with a circular cross section, increasing the accelerator system area is accomplished by increasing the engine diameter. This has led to the development of engines sizes ranging from 5 to 150 cm in diameter over the past 30 years.
To maintain a constant thrust (and power) density as the engine diameter is increased requires that the grid-to-grid separation remain constant. This requirement results in increasing values of the grid span-to-gap ratio, i.e., the ratio of accelerator system diameter to the grid separation. The current state-of-the-art 30-cm diameter ion accelerator system has a span-to-gap ratio of approximately 500. The maximum achievable span-to-gap ratio is limited by mechanical constraints imposed by fabrication and handling procedures, as well as by thermal effects which serve to alter the grid separation during engine operation.
A conventional circular ion engine using argon propellant and operating at a specific impulse of 10,000 seconds would require a beam diameter of approximately 2.2 m to process one megawatt. Assuming a maximum electric field between the grids of 3000 V/mm, this thruster would require the development of an accelerator system with a span-to-gap ratio of about 1700. This is a factor of 3.4 beyond the state of the art, and would have to be developed for an engine diameter which is more than a factor of seven greater than the present 30-cm thruster.
Aside from increasing the active grid area, the power processed by an ion engine may be increased by increasing the net accelerating voltage. For a given propellant this voltage determines the engine specific impulse. For the Mars cargo and piloted Mars missions using electric propulsion, specific impulses in the range 7,000 to 10,000 seconds are required. With argon propellant, this translates into net accelerating voltages which are roughly a factor of two higher than that typically used on the 30-cm thruster with xenon propellant (which was designed for operation at specific impulses less than 4,000 seconds).
Finally, the use of lighter atomic mass propellants increases the current handling capability of the accelerator system at a given voltage, which in turn increases the power processed by the engine. Therefore, to scale ion engines up to megawatt power levels it is necessary to significantly increase the active accelerator system area, operate at high applied net accelerating voltages, and use light atomic mass propellants. The last two of these items must together be consistent with the specific impulse range required for the application of the high-power engine.
The development of 100-kW and megawatt class ion engines must be achieved primarily by scaling up the active grid area for beam extraction by one or two orders of magnitude from the current state of the art. To overcome span-to-gap limitations associated with continuously increasing the diameter of the conventional circular ion engine, alternate engine geometries have been proposed, including an annular engine configuration (such as described in the paper by Aston, G. and Brophy, J. R., "A 50-cm Diameter Annular Ion Engine," AIAA-89-2716, July 1989) and a rectangular engine design (such as described in the paper by Gilland, J. H., Myers, R. M. and Patterson, M. J., "Multimegawatt Electric Propulsion System Design Considerations," AIAA-90-2552, July 1990).